1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine vane with a heat shield on the shroud.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section with multiple stages of stator vanes and rotor blades to extract mechanical energy from a hot gas flow passing from the combustor and through the turbine. Stator vanes guide the gas flow into the rotor blades for higher efficiency. The stator vanes and rotor blades include complex internal cooling passages and film cooling hole arrangements to provide cooling of the airfoils in order that a higher temperature can be used in the turbine. Higher temperatures result in higher efficiencies.
The stator vanes are located upstream of an adjacent rotor blade arrangement. The stator vanes include an airfoil portion that extends between an inner and an outer shroud. The inner and outer shrouds form a flow guiding surface that is exposed to the hot gas flow. The shrouds are also cooled by passing cooling air along the inner surface and with film cooling holes that supply a jet of film cooling air into the hot gas flow. FIG. 1 shows a prior art turbine vane endwall leading edge region that is cooled with a double row of circular or shaped film cooling holes. In the FIG. 1 vane, a streamwise and circumferential cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. Film cooling air that is discharged from the double film rows have a tendency to migrate from the pressure side toward the vane suction surface which induces a mal-distribution of film cooling flow and endwall metal temperature. Multiple rows of shaped discrete film holes are used for this cooling of the pressure side and suction side of the endwall surfaces. As a result of this cooling approach, a large amount of cooling air is used for the cooling of vane endwall surface which yields a high mixing loss for the turbine stage due to cooling air interacting with the mainstream hot gas flow. The mixing losses are especially higher for the cooling rows that discharge beyond the gage point.
It is an object of the present invention to provide for a turbine stator vane with better cooling for the inner and outer shrouds.
It is another object of the present invention to provide for better cooling of the inner and outer shrouds of the turbine stator vanes which make use of less cooling air.
It is another object of the present invention to provide cooling for the turbine stator vane shrouds which eliminate the use of active film cooling holes for the vane endwall and therefore greatly reduce the mixing loses due to cooling air interaction with the main stream hot gas flow.